JP3947227B2 - Gas turbine engine shroud seal - Google Patents

Gas turbine engine shroud seal Download PDF

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JP3947227B2
JP3947227B2 JP54130897A JP54130897A JP3947227B2 JP 3947227 B2 JP3947227 B2 JP 3947227B2 JP 54130897 A JP54130897 A JP 54130897A JP 54130897 A JP54130897 A JP 54130897A JP 3947227 B2 JP3947227 B2 JP 3947227B2
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Prior art keywords
shroud
seal
groove
radial
platform
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JP2000511256A (en
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ピジー，アントニオ．
シー． クローネ，ジェイムス
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Pratt and Whitney Canada Corp
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Pratt and Whitney Canada Corp
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F16—ENGINEERING ELEMENTS AND UNITS; GENERAL MEASURES FOR PRODUCING AND MAINTAINING EFFECTIVE FUNCTIONING OF MACHINES OR INSTALLATIONS; THERMAL INSULATION IN GENERAL
    • F16J—PISTONS; CYLINDERS; SEALINGS
    • F16J15/00—Sealings
    • F16J15/02—Sealings between relatively-stationary surfaces
    • F16J15/06—Sealings between relatively-stationary surfaces with solid packing compressed between sealing surfaces
    • F16J15/08—Sealings between relatively-stationary surfaces with solid packing compressed between sealing surfaces with exclusively metal packing
    • F16J15/0887—Sealings between relatively-stationary surfaces with solid packing compressed between sealing surfaces with exclusively metal packing the sealing effect being obtained by elastic deformation of the packing
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/005—Sealing means between non relatively rotating elements
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/24—Casings; Casing parts, e.g. diaphragms, casing fastenings
    • F01D25/246—Fastening of diaphragms or stator-rings
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/11—Shroud seal segments

Description

背景技術
１．発明の分野
本発明は、ガスタービンエンジンに関し、より詳細にはガスタービンエンジンを取り囲むケーシング内において冷却流体の漏れを低減させるためのタービンブレード用のシュラウドシールに関する。
２．従来技術
タービンブレードの先端部を取り囲む環状シュラウドのシール性能を改善するため、多くの試みがなされてきている。エンジンケーシングは、航空機エンジンのタービン領域を取り囲んでおり、ケーシング内の種々の要素例えば、シュラウド自体等に冷却ガスを送っている。エンジンのタービン領域は、ケーシングと同心とされており、極めて高温のガスの通路となっている。このように取り囲むケーシング構造体は、相対的に低温に維持することが好ましい。
タービンブレード先端部のための環状のシュラウドは、通常周方向にセグメント化されている。シールは、シュラウドセグメントの間におけるギャップでのガスリークを防止するために必要となる。クレベンガー（Clevenger）等の１９８８年８月３０日に付与された米国特許第4,767,260号及びディクソン（Dixon）等に１９９２年１０月２７日に付与された米国特許第5,158,430号では、ベーンアッセンブリのプラットホームの間に配設された種々のフェザーシールが開示されている。サンディジュニア（Sandy, Jr）等に１９８６年３月４日に付与された米国特許第4,573,866号は、ブレード先端シュラウドの端部の間における径方向平面に用いられるフェザーシールを開示している。ベイレイ（Bailey）等に１９９４年６月７日に付与された米国特許第5,318,402号では、スプライン又はフェザーシール及びシュラウドとフェザーシールとをロックするための周方向のスペーサが開示されている。
異なったシーリングリングは、また異なったシュラウド及びケーシング配置において必要とされる。例えば、分離した連結部が軸方向の構成要素を有している溝を備えた形状とされている場合には、“Ｃ”形状のリングとされるのが通常である。米国特許第4,573,866号においては、ベローズタイプリングがこのような溝に用いられている。このようなリングは、良好なシーリング特性を与えるが、それらの径方向の断面の故に圧縮に対して固すぎ、したがってエンジンの保守の際に容易に交換できない。実際、上述のシールを交換する場合には、特殊工具が必要とされていた。
発明の開示
本発明は、軸方向構成要素と径方向構成要素とを備えたタイプの改善されたシュラウドセグメント間シールを提供することにある。
本発明はさらに、高温ガス流路に冷却ガスがリークするのを低減するための周方向のシュラウドセグメント端部の間に形成されるギャップをシールするためのワンピースシートシールを提供することを目的とする。
本発明はさらに、軸方向構成要素を備えた環状ギャップの形成された環状ジョイント部をシールするためのリングシールを提供することを目的とする。
本発明は、さらにより簡単に装着するようにするために、径方向に容易に圧縮できるシートメタルから形成されたリングシールを提供することを目的とする。
本発明はさらに、タービンブレード先端シュラウドアッセンブリのための改善されたシール構成を提供することを目的とする。
本発明の構成によれば、シュラウドがガスタービンエンジンの軸方向に沿った高温ガス流路と同心の冷却ガス通路の間に配設され、前記冷却ガス通路は、前記シュラウドと外側ケーシングとの間に形成され、シュラウドが周方向に配列されたセグメントから構成され、各シュラウドセグメントが軸方向のプラットホームとこのプラットホームから延びた少なくとも１つの径方向リブとを有しており、前記シュラウドセグメントは端部壁とこの端部壁に形成され、プラットホーム及び前記リブに対応する端部壁位置に形成された連続溝とを有しており、シールは、端部壁部分の連続溝の形状にされたワンピースシート部材からなり、シートの一部として、軸方向構成要素と径方向構成要素とを有するガスタービンエンジンのタービンブレード先端シュラウドのためのシールが提供される。
本発明のより詳細な特定の実施例においては、シールを形成するシートの径方向構成要素は、２重とされており、この径方向構成要素が、径方向リブの壁に形成される溝と摩擦力を生じさせるための弾性部材となる。
本発明の別の特徴によれば、ガスタービンエンジンにはリングシールが装着され、環状のギャップが軸方向構成要素に形成された場合に、２つの環状エンジン部品の間を連結させていて、この軸方向のギャップは、互いに離間した径方向壁を有し、リングシールが波打った部分を備える構成とされ、この波打った部分は、一つの波打ち部分のピークを有しているとともに、ギャップの離間した壁のうちの一方に接触し、他方のピークがギャップの他の離間した壁と接触していて、このギャップからジョイント部の両側へとガスが漏れて行くのを防止しているとともに、波打ち部分は、充分にシールが径方向へと容易に圧縮されるような角度とされている。
本発明のより具体的な実施例によれば、ガスタービンエンジンのタービンブレード先端部シュラウドにはリングシールが配設されており、シュラウドは、軸方向に向いた高温ガス通路と、該シュラウドと外側ケーシングの間に形成される冷却ガス通路と、の間に配置されている。シュラウドは、軸方向に延びたプラットホームを有しているとともに、少なくとも径方向に延びたリブを備えていて、このリブがケーシング上の取り付け手段に連結されるようになっている。環状のギャップは、シュラウドとケーシングの取り付け手段の間に形成された軸方向の構成要素を有しており、環状ギャップは、シュラウドと取り付け手段とに形成された径方向に互いに離間した壁を有し、前記リングシールは、“Ｃ”形状の脚部部分が軸方向に延びていて、この“Ｃ”部分がその軸方向の脚部よりも大きな径方向の寸法を有し、かつ、この“Ｃ”部分は、ギャップの半径方向部分内に配置されかつ、ギャップの径方向寸法よりも小さな径方向寸法を有した“Ｃ”形状を有し、リングシールの脚部は、交互にピークを形成する波打ちパターンを備えていて、このピークは、環状ギャップの径方向に互いに離間した壁に接触して、シュラウドと取り付け手段の間のジョイント部分におけるガスシールを形成して、冷却ガスフローの漏れを防止するとともに、前記波打ちパターンは、シールを径方向の圧縮が容易になるような角度とされている。
【図面の簡単な説明】
本発明の特徴について、本発明の好適な実施例を図示した添付の図面を参照して説明する。
図１は、本発明の１つの実施例の一部分解して示した部分斜視図であり；
図２は、図１の詳細な正面図であり；
図３は、本発明の別実施例を示した一部断面斜視図であり；
図４は、本発明の詳細部を示したガスタービンエンジンのタービン領域の軸方向断面図であり；
図５は、図４と同一の平面に沿った拡大部分断面図であり、図４の詳細部を示した図である。
好適な実施例の説明
図面を参照すると、特に図１及び図２には、エンジン１０に用いられる典型的なタービンブレードシュラウド１２が示されている。シュラウド１２は、複数のセグメント１２ａ．．．１２ｎを有しており、これらのセグメントは、周方向に配列されているとともに、タービンブレード３８がマウントされるロータと同心とされている（図１及び図４）。
シュラウドセグメントは、図１においては参照符号１２により示されていて、軸方向構成要素を有する本質的に環状のプレート様部材として形成されたプラットホーム１４と、それぞれフランジ２０，２２を備え、対となった上側に延びたリブ１６，１８とを有している。これらのリブ１６，１８及びそれぞれのフランジ２０及び２２は、冷却空気の通路及びチャンバを画成する他、シュラウドプラットホーム１４の支持体としても機能する。フランジ２０，２２は、図４に示されているようにまた、エンジンケーシング３２内にシュラウドを取り付けているが、この取り付け構造体３６については、後述する。エンジンケーシング３２の壁３４には、換気用開口３３が設けられていて、支持構造体とケーシング３２の間の空間に冷却空気が流入するようにさせている。換気用開口３１は、また、支持構造体３６に設けられていて、この冷却空気をシュラウド１２と支持構造体３６の間に形成された環状チャンネル３５へと流入するようにさせている。最後に、シュラウド１２に形成されたアパーチャ３７は、冷却空気が高温ガス流路に排気されるようにしている。
セグメント間シールを提供するために、各シュラウドセグメント１２の端部壁には、連続溝２６が配設されており、この連続溝は、軸方向構成要素２６ｂと、リブ１６，１８に対応する径方向構成要素２６ａ，２６ｃを有している。シール２４は、耐熱性合金のシート２４から製造されていて、このシートが、立ち上がった径方向構成要２８，３０とプラットホーム１４に対応する水平方向部分つまり軸方向構成要素２５，２７とを形成するように曲げられている。このシールは、当然ながらいかなる好適な耐熱性材料から押し出し成型又はモールドすることもできる。
立ち上がった径方向構成要素２８，３０は、それぞれ脚部２８ａ，２８ｂ，３０ａ，３０ｂを有していて、これらの脚部は、僅かに広げられてそれぞれの溝よりも広くされており、これが図２に示されている。この構成は、径方向構成要素２８，３０に対して弾性の度合いを与えており、シール２４がシュラウド１２の端部における溝２６に嵌合された場合に、径方向構成要素２８，３０は、それぞれ上向きの溝セグメント２６ａ，２６ｃに挿入されて圧縮され、溝２６にシールを堅固に固定するようになっている。その弾性の観点から、セグメントは容易に組み立てることができ、その後分離することもでき、保守が必要とされた場合にはシール２４は容易に交換することができる。
本発明の別の特徴は、図３及び図５に示されている。エンジン１０のタービン領域は、図４にその一部が示されていて、これに加えてタービンブレード３８と、ステータベーン４２，４４とが示されている。各ステータベーンは、ステータベーンに取り付けられたシュラウド構造体を有しており、シュラウド１２は、タービンブレード３８のブレード先端部４０を取り囲んでいる。
図に示されているように、エンジンケーシング３２内の取り付け構造体３６は、シュラウド１２を支持しており、これが図４及び図５に示されている。シュラウド１２のフランジ２２は、例えば支持構造体３６のフランジ４７に支持されていても良い。同様にして、フランジ２０は、支持構造体３６のフランジによって支持されていても良い。リブ１６，１８は、支持構造体３６の構造とともに、シュラウドの内側ならびにこれを取り囲んだ冷却チャンネル６４，６６を形成している。冷却チャンネル６４内の空気の圧力は、周囲のチャンネル６６とは異なっており、したがってこれら異なったチャンネルすなわちチャンバは、シールをする必要がある。
リングシール４６は、フランジ４７によって形成された溝により画成されたギャップ、すなわち環状の溝４８に設けられている。この場合、溝４８は、フランジ２２が挿入された状態では、軸方向構成要素４８ａと径方向構成要素４８ｂを有する“Ｌ”形状をなす。リングシール４６は、“Ｃ”形状の部分５２を有しており、この“Ｃ”形状部分５２は、径方向構成要素４８ｂの領域における溝４８よりも径寸法が小さくされている。“Ｃ”形状部分５２は、脚部５４，５６を有している。脚部５４の延長部は、フラットな角度に面した波打ちパターンを有していて、ピーク６０，６２を備えた波打ち領域５８を構成している。アーム５８ａ，５８ｂによって規定される角度は、鈍角とされている。この角度は鋭角とすることもできるが、シールが容易に径方向の圧縮を行うことができるだけ充分に大きくされている。
波打ち領域５８の形状は、ピーク６０，６２の間の径方向の寸法が、溝４８の壁３６ａ，２２ａの間の径方向寸法よりも僅かに大きくなるようにされている。リングシール４６は、対向した壁２２ａ，３６ａに対して継続的にスプリングによるフィッティングを維持させるように、本質的に弾性を有する耐熱性材料又は耐熱性合金から形成されている。したがって、波打ち領域５８は、溝４８に対してスプリングフィットされているので、ピーク６０，６２は、対向する壁２２ａ，３６ａに対して継続的に接触している。
したがって、リングシール４６の形状により、このリングシールは、組立の際に、シュラウドフランジ２２が内部に挿入される前に、溝４８にフィットさせることができる。リングシール４６は、溝４８にフィットさせることが可能なエンドレスリングとして構成することもできる。
フランジ２２は、溝４８内にすで取り付けられているスプリングシール４６に対して挿入できるようにするために、図示されているように傾斜部を有していても良い。波打ち部分５８の径方向における弾性は、保守の間にシュラウドが溝４８から取り外される際にリングシールを容易に交換可能とさせている他、フランジ２２の挿入を容易にさせている。
したがって、シール２４とリングシール４６は、双方ともシュラウド１２と協働してシール及びシュラウドセグメントのアッセンブリへの装着を容易にさせているとともに、シュラウド１２の保守を著しく容易とさせている。Background Art The present invention relates to gas turbine engines, and more particularly to turbine blade shroud seals for reducing cooling fluid leakage within a casing surrounding the gas turbine engine.
2. Many attempts have been made to improve the sealing performance of an annular shroud surrounding the tip of a prior art turbine blade. The engine casing surrounds the turbine area of the aircraft engine and delivers cooling gas to various elements within the casing, such as the shroud itself. The turbine region of the engine is concentric with the casing and is a very hot gas passage. The surrounding casing structure is preferably maintained at a relatively low temperature.
An annular shroud for the turbine blade tip is typically segmented in the circumferential direction. Seals are required to prevent gas leaks in the gap between the shroud segments. U.S. Pat. No. 4,767,260 granted Aug. 30, 1988 to Clevenger et al. And U.S. Pat. No. 5,158,430 granted Oct. 27, 1992 to Dixon et al. Various feather seals disposed between are disclosed. U.S. Pat. No. 4,573,866, issued Mar. 4, 1986 to Sandy Jr. et al., Discloses a feather seal for use in a radial plane between the ends of a blade tip shroud. U.S. Pat. No. 5,318,402 issued Jun. 1994 to Bailey et al. Discloses a spline or feather seal and a circumferential spacer for locking the shroud and feather seal.
Different sealing rings are also required in different shroud and casing arrangements. For example, when the separated connecting portion has a shape including a groove having an axial component, it is usually a “C” shaped ring. In US Pat. No. 4,573,866, a bellows type ring is used for such a groove. Such rings provide good sealing properties but are too hard to compress because of their radial cross-section and therefore cannot be easily replaced during engine maintenance. In fact, a special tool was required to replace the above-described seal.
Disclosure of the Invention It is an object of the present invention to provide an improved inter-shroud segment seal of the type comprising an axial component and a radial component.
The present invention further aims to provide a one-piece seat seal for sealing a gap formed between circumferential shroud segment ends to reduce leakage of cooling gas into the hot gas flow path. To do.
It is a further object of the present invention to provide a ring seal for sealing an annular joint portion having an annular gap with an axial component.
It is an object of the present invention to provide a ring seal formed from sheet metal that can be easily compressed in the radial direction in order to make it even easier to install.
The present invention further seeks to provide an improved seal arrangement for a turbine blade tip shroud assembly.
According to the configuration of the present invention, the shroud is disposed between the hot gas passage along the axial direction of the gas turbine engine and the concentric cooling gas passage, and the cooling gas passage is between the shroud and the outer casing. Each of the shroud segments includes an axial platform and at least one radial rib extending from the platform, the shroud segment having an end portion. One piece formed in the end wall portion and having a continuous groove formed at the end wall position corresponding to the platform and the rib, the seal being formed in the shape of a continuous groove in the end wall portion. A turbine blade tip of a gas turbine engine comprising a seat member and having an axial component and a radial component as part of the seat Seal for the shroud is provided.
In a more specific specific embodiment of the invention, the radial component of the sheet forming the seal is doubled, and this radial component is a groove formed in the wall of the radial rib. It becomes an elastic member for generating a frictional force.
According to another feature of the present invention, when the gas turbine engine is fitted with a ring seal and an annular gap is formed in the axial component, there is a connection between the two annular engine components. The axial gap has radial walls that are spaced apart from each other, and the ring seal is configured to have a corrugated portion, the corrugated portion having a peak of one corrugated portion and a gap. One of the spaced walls and the other peak is in contact with the other spaced walls of the gap, preventing gas from leaking from the gap to both sides of the joint. The wavy portion is sufficiently angled so that the seal is easily compressed in the radial direction.
In accordance with a more specific embodiment of the present invention, a ring seal is disposed on a turbine blade tip shroud of a gas turbine engine, the shroud comprising an axially oriented hot gas passage, the shroud and an outer side. And a cooling gas passage formed between the casings. The shroud has an axially extending platform and is provided with at least a radially extending rib that is connected to mounting means on the casing. The annular gap has an axial component formed between the shroud and the attachment means of the casing, and the annular gap has radially spaced walls formed in the shroud and the attachment means. The ring seal has a "C" shaped leg portion extending in the axial direction, the "C" portion having a larger radial dimension than the axial leg portion, and the "C" portion. The “C” portion has a “C” shape that is located within the radial portion of the gap and has a radial dimension that is smaller than the radial dimension of the gap, and the legs of the ring seal alternately peak. This peak is in contact with the radially spaced walls of the annular gap, forming a gas seal at the joint between the shroud and the mounting means, and the cooling gas flow. Thereby preventing leakage of the wavy pattern is an angle such that seal radially compressive becomes easy.
[Brief description of the drawings]
The features of the present invention will be described with reference to the accompanying drawings, which illustrate preferred embodiments of the present invention.
FIG. 1 is a partial perspective view, partly exploded, of one embodiment of the present invention;
FIG. 2 is a detailed front view of FIG. 1;
FIG. 3 is a partial sectional perspective view showing another embodiment of the present invention;
4 is an axial cross-sectional view of the turbine region of a gas turbine engine showing details of the present invention;
FIG. 5 is an enlarged partial cross-sectional view along the same plane as FIG. 4 and shows a detailed portion of FIG.
DESCRIPTION OF THE PREFERRED EMBODIMENTS Referring to the drawings, a typical turbine blade shroud 12 for use in an engine 10 is shown in particular in FIGS. The shroud 12 includes a plurality of segments 12a. . . 12n, these segments are arranged circumferentially and are concentric with the rotor on which the turbine blades 38 are mounted (FIGS. 1 and 4).
The shroud segment, indicated in FIG. 1 by reference numeral 12, comprises a platform 14 formed as an essentially annular plate-like member having axial components and flanges 20 and 22, respectively, in pairs. And ribs 16 and 18 extending upward. These ribs 16 and 18 and the respective flanges 20 and 22 define a passage and chamber for cooling air, and also function as a support for the shroud platform 14. As shown in FIG. 4, the flanges 20 and 22 also have a shroud attached in the engine casing 32. The attachment structure 36 will be described later. The wall 34 of the engine casing 32 is provided with a ventilation opening 33 so that cooling air flows into the space between the support structure and the casing 32. A ventilation opening 31 is also provided in the support structure 36 to allow this cooling air to flow into an annular channel 35 formed between the shroud 12 and the support structure 36. Finally, the aperture 37 formed in the shroud 12 allows the cooling air to be exhausted to the hot gas flow path.
In order to provide an inter-segment seal, an end wall of each shroud segment 12 is provided with a continuous groove 26 that has a diameter corresponding to the axial component 26b and the ribs 16,18. It has direction component 26a, 26c. The seal 24 is manufactured from a heat-resistant alloy sheet 24 that forms a raised radial component 28, 30 and a horizontal portion corresponding to the platform 14, ie, axial components 25, 27. Is bent like so. The seal can of course be extruded or molded from any suitable refractory material.
The raised radial components 28, 30 have legs 28a, 28b, 30a, 30b, respectively, which are slightly widened and wider than the respective grooves. 2. This configuration provides a degree of elasticity to the radial components 28, 30, and when the seal 24 is fitted in the groove 26 at the end of the shroud 12, the radial components 28, 30 are Each is inserted into an upward groove segment 26 a, 26 c and compressed to firmly fix the seal in the groove 26. Because of its elasticity, the segments can be easily assembled and then separated, and the seal 24 can be easily replaced if maintenance is required.
Another feature of the present invention is illustrated in FIGS. A part of the turbine region of the engine 10 is shown in FIG. 4, and in addition, a turbine blade 38 and stator vanes 42 and 44 are shown. Each stator vane has a shroud structure attached to the stator vane, and the shroud 12 surrounds the blade tip 40 of the turbine blade 38.
As shown, the mounting structure 36 in the engine casing 32 supports the shroud 12, which is shown in FIGS. 4 and 5. The flange 22 of the shroud 12 may be supported by the flange 47 of the support structure 36, for example. Similarly, the flange 20 may be supported by the flange of the support structure 36. The ribs 16 and 18 together with the structure of the support structure 36 form cooling channels 64 and 66 that surround the shroud and surround it. The pressure of the air in the cooling channel 64 is different from the surrounding channel 66, so these different channels or chambers need to be sealed.
The ring seal 46 is provided in a gap defined by a groove formed by the flange 47, that is, an annular groove 48. In this case, the groove 48 has an “L” shape having an axial component 48a and a radial component 48b when the flange 22 is inserted. The ring seal 46 has a “C” shaped portion 52 that is smaller in diameter than the groove 48 in the region of the radial component 48b. The “C” shaped portion 52 has legs 54 and 56. The extension part of the leg part 54 has a corrugated pattern facing a flat angle, and constitutes a corrugated region 58 having peaks 60 and 62. The angle defined by the arms 58a and 58b is an obtuse angle. This angle can be an acute angle, but is made large enough to allow the seal to easily compress in the radial direction.
The shape of the wavy region 58 is such that the radial dimension between the peaks 60 and 62 is slightly larger than the radial dimension between the walls 36a and 22a of the groove 48. The ring seal 46 is made of a heat-resistant material or heat-resistant alloy having essentially elasticity so that the fitting by the spring is continuously maintained with respect to the opposed walls 22a and 36a. Accordingly, since the wavy region 58 is spring-fitted with respect to the groove 48, the peaks 60 and 62 are in continuous contact with the opposing walls 22a and 36a.
Thus, the shape of the ring seal 46 allows the ring seal to fit into the groove 48 during assembly before the shroud flange 22 is inserted therein. The ring seal 46 can also be configured as an endless ring that can fit into the groove 48.
The flange 22 may have a ramp as shown to allow insertion into a spring seal 46 that is already installed in the groove 48. The radial elasticity of the undulating portion 58 makes it easy to replace the ring seal when the shroud is removed from the groove 48 during maintenance and facilitates the insertion of the flange 22.
Thus, both the seal 24 and the ring seal 46 cooperate with the shroud 12 to facilitate the mounting of the seal and shroud segment to the assembly and greatly facilitate the maintenance of the shroud 12.

Claims (3)

軸方向に向いた高温ガス通路と同心の冷却ガス通路との間にタービンブレード先端シュラウド（１２）が配置され、前記冷却ガス通路が、前記シュラウド（１２）と外側ケーシング（３２）との間に形成されたガスタービンエンジンのためのタービンブレード先端シュラウドシール（２４）であって、
前記シュラウド（１２）は、周方向に配列されたセグメント（１２ａ．．．１２ｎ）を有しており、
前記各シュラウドセグメント（１２ａ．．．１２ｎ）は、軸方向のプラットホーム（１４）と該プラットホーム（１４）から延びた少なくとも１つの径方向リブ（１６、１８）を有しており、
前記シュラウドセグメント（１２ａ．．．１２ｎ）は、端部壁を備えるとともに、前記プラットホーム（１４）ならびに前記リブ（１６、１８）に対応して各端部壁内に形成された連続する溝（２６）を備えており、
前記シール（２４）は、耐熱性材料からなるワンピースのシート材料により前記端部壁の連続する溝（２６）の形状に形成されているとともに、径方向構成要素（２８、３０）と軸方向構成要素（２５、２７）とを有し、
前記径方向構成要素（２８、３０）は、前記端部壁の溝（２６）の中の前記径方向リブ（１６，１８）に対応する部分の溝（２６ａ、２６ｃ）よりも大きな幅を有するように前記シート材料を２重に折り返してなり、
前記軸方向構成要素（２５、２７）は、前記端部壁の溝（２６）の中の前記軸方向プラットホーム（１４）に対応する部分の溝（２６ｂ）に嵌合する、ことを特徴とするシール（２４）。A turbine blade tip shroud (12) is disposed between the axially oriented hot gas passage and the concentric cooling gas passage, the cooling gas passage between the shroud (12) and the outer casing (32). A turbine blade tip shroud seal (24) for a formed gas turbine engine comprising:
The shroud (12) has circumferentially arranged segments (12a ... 12n),
Each shroud segment (12a ... 12n) has an axial platform (14) and at least one radial rib (16, 18) extending from the platform (14);
The shroud segments (12a... 12n) include end walls and continuous grooves (26) formed in each end wall corresponding to the platform (14) and the ribs (16, 18). )
The seal (24) is formed in the shape of a continuous groove (26) in the end wall by a one-piece sheet material made of a heat-resistant material, and has a radial component (28, 30) and an axial configuration. Elements (25, 27),
The radial components (28, 30) have a larger width than the portion of the grooves (26a, 26c) corresponding to the radial ribs (16, 18) in the groove (26) of the end wall. The sheet material is folded twice as follows,
The axial component (25, 27) is fitted in a groove (26b) in a portion corresponding to the axial platform (14) in the end wall groove (26). Seal (24). 前記材料は、所望する形状に曲げることができる耐熱合金から構成されていることを特徴とする請求項１に記載のシール（２４）。The seal (24) of claim 1, wherein the material comprises a heat resistant alloy that can be bent into a desired shape. 前記シュラウド（１２）は、互いに離間して前記プラットホーム（１４）から半径方向に延びた一対の平行リブ（１６、１８）を有するとともに、前記溝（２６）は双方のリブ（１６、１８）内に連続して延びており、
前記シール（２４）は、実質的に同一形状の互いに離間した一対の径方向構成要素（２８、３０）を有し、各々の径方向構成要素（２８、３０）が、前記シュラウドセグメント（１２ａ．．．．１２ｎ）の端部壁の前記連続溝（２６）における前記リブ（１６、１８）に対応する部分（２６ａ、２６ｃ）に弾性的に嵌合することを特徴とする請求項１に記載のシール（２４）。The shroud (12) has a pair of parallel ribs (16, 18) spaced from each other and extending radially from the platform (14), and the groove (26) is in both ribs (16, 18). Extending continuously,
The seal (24) has a pair of spaced apart radial components (28, 30) of substantially the same shape, each radial component (28, 30) being connected to the shroud segment (12a. 2.. 12 n) end-walls of the continuous groove (26) corresponding to the ribs (16, 18) elastically fitting into portions (26a, 26c). Seal (24).
JP54130897A 1996-05-20 1997-05-20 Gas turbine engine shroud seal Expired - Fee Related JP3947227B2 (en)

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